1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine blade with leading edge film cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
The first and second stage turbine stator vanes and rotor blades are exposed to the highest gas flow temperatures in the turbine, and therefore require film cooling of the external surfaces. The leading edge region of these airfoils is exposed to the highest gas flow temperatures on the airfoil. To provide film cooling for the leading edge region, a showerhead arrangement of film cooling holes is used. FIG. 1 shows a prior art blade with a cooling circuit that includes three rows of film cooling holes in a showerhead arrangement on the leading edge region with two rows of gill holes on both sides therefore. FIG. 2 shows a detailed view of the leading edge region with the film holes and gill holes. The middle row of film cooling holes is located at a stagnation line of the leading edge which is the location of the highest heat load on the airfoil. Cooling air from a supply channel 11 is metered through a row of metering holes 12 and into a leading edge impingement cavity 13 from which the showerhead film cooling holes 14 are supplied with the cooling air.
FIG. 3 shows a cross section view through the middle row of film cooling holes which are angled at from 20 to 35 degrees from the airfoil surface. The showerhead arrangement of film cooling holes in FIGS. 1-4 suffer from several disadvantages. The heat load on the blade leading edge region is parallel to the film cooling hole arrangement and thus reduces the cooling effectiveness. A portion of the film cooling holes within each film row is positioned behind each other (see FIG. 3) which reduces the effective frontal convection cooling area and conduction distance for the oncoming heat load. A realistic minimum film hole spacing to diameter ratio is around 3.0 such that below this ratio and cracking may occur for the film row. This results in a maximum achievable film coverage for that particular film row of 33% or 0.33 film effectiveness for each row of film cooling holes in the showerhead arrangement. Since the showerhead film cooling holes are at a radial orientation, the film pattern discharged from the film holes overlap with each other as depicted in FIG. 4. The film layer from the middle row flows over the film holes on the two outer rows and thus leaves a space 15 between film holes in the row that is uncovered. This is especially the case for a rotor blade because of the rotational effects on the film discharge.